1. Field of the Invention
The invention relates generally to aircraft fuselage design. More particularly, it relates to aircraft fuselage configurations that control the build-up, magnitude, and directional characteristics of pressure waves generated by an aircraft flying at supersonic speed so as to reduce the sonic boom at ground level.
2. The Prior Art and Background
A conventional aircraft in flight produces pressure waves in the medium it flies through. These pressure waves propagate at the speed of sound. When the aircraft flies at subsonic speed, these pressure waves propagate in all directions around the aircraft, including ahead of the aircraft. When the aircraft flies at supersonic speed, these pressure waves cannot propagate ahead of the aircraft because the aircraft is traveling faster than the
propagation speed of the waves. Instead, the pressure waves generated by such a conventional aircraft coalesce into two shock waves, one formed by the nose of the aircraft and the other formed by the tail of the aircraft. These shock waves are characterized by an abrupt pressure increase across the wave. With respect to the shock wave formed by the nose (the xe2x80x9cbow shockxe2x80x9d), the pressure increases abruptly from about ambient to above ambient. The pressure decreases from above ambient to below ambient in the region between the bow shock and the shock wave formed by the tail (the xe2x80x9ctail shockxe2x80x9d). The pressure then increases abruptly from below ambient to about ambient across the tail shock.
In the limit, weak shock waves propagate in the form of a Mach cone having a shape defined by the Mach angle xcexc. The Mach angle xcexc is a function of the Mach number M, which is defined as the ratio of the speed of an object to the speed of sound. The Mach angle xcexc can be determined using the equation:                     sin        ⁡                  (          μ          )                    =              1        M              ,    or        μ    =                  sin                  -          1                    ⁡              (                  1          M                )            
The shape of the Mach cone produced by an aircraft in supersonic flight can be represented by rotating a line drawn at an angle xcexc to the aircraft""s direction of travel about a line representing the aircraft""s direction of travel, so that the tip of the Mach cone points in the direction of travel.
These shock waves can propagate great distances away from the aircraft and eventually reach the ground, where they can produce significant acoustic disturbances called sonic booms. Sonic booms are so named because of the sounds created by the abrupt pressure changes when the shock waves pass a reference point on the ground. The acoustic signature of a sonic boom commonly is characterized as an N-wave because the pressure changes associated with the acoustic signature resemble the letter xe2x80x9cNxe2x80x9d when plotted as a function of aircraft length. That is, an N-wave is characterized by the abrupt pressure rise associated with the bow shock, commonly referred to as xe2x80x9cpeak overpressure,xe2x80x9d followed by a substantially linear decrease in pressure to below ambient pressure, followed by the abrupt rise to ambient pressure associated with the tail shock. Sonic booms can cause objectionable sounds and vibrations. For these reasons, supersonic flight over populated areas has long been limited by regulation.
In order for supersonic flight over land to be acceptable, the pressure disturbances that cause the sonic boom""s acoustic signature must be controlled so that the effects of the abrupt pressure changes caused by the shock waves are minimized at ground level. Many attempts have been made to modify the design of supersonic aircraft in order to adjust the sonic boom signature. These modifications have included changes to wing design, as described in U.S. Pat. No. 5,934,607, issued to Rising, et al., for a xe2x80x9cShock Suppression Supersonic Aircraft.xe2x80x9d Another approach involves incorporating air passages through the fuselage or wings of supersonic aircraft, such as the structures described in U.S. Pat. No. 4,114,836, issued to Graham, et al., for an xe2x80x9cAirplane Configuration Design for the Simultaneous Reduction of Drag and Sonic Boomxe2x80x9d; U.S. Pat. No. 3,794,274, issued to Eknes, for an xe2x80x9cAircraft Structure to Reduce Sonic Boom Intensityxe2x80x9d; and U.S. Pat. No. 3,776,489, issued to Wen, et al., for a xe2x80x9cSonic Boom Eliminator.xe2x80x9d Further attempts at reducing the sonic boom caused by supersonic aircraft include the addition to the aircraft of structure arranged to disrupt the air flow patterns as the aircraft travels at supersonic speed. Examples include the structure described in U.S. Pat. No. 3,709,446, issued to Espy, for a xe2x80x9cSonic Boom Reductionxe2x80x9d and U.S. Pat. No. 3,647,160, issued to Alperin, for a xe2x80x9cMethod and Apparatus for Reducing Sonic Booms.xe2x80x9d
Another attempt to control the sonic boom in a supersonic aircraft uses a blunt nose to increase the air pressure immediately adjacent to the nose of the aircraft, thus disrupting the normal formation of the pressure wave that causes the acoustic signature. This disruption results in the reduction of abrupt pressure changes in the acoustic wave that strikes the ground. A blunt nose, however, creates a significant amount of drag on the aircraft, drastically decreasing its efficiency. U.S. Pat. No. 5,740,984, issued to Morgenstern, for a xe2x80x9cLow Sonic Boom Shock Control/Alleviation Surfacesxe2x80x9d describes a mechanical device on the nose of the airplane which can be moved between a first position effecting a blunt nose when sonic boom reduction is desired and a second position effecting a streamlined nose when sonic boom reduction is not required, thereby removing (in the streamlined configuration) the drag penalty inherent in a blunt nose design. U.S. Pat. No. 4,650,139, issued to Taylor et al., discloses a blunt-nosed spike which can be extended from a space vehicle""s fuselage.
U.S. Pat. Nos. 5,358,156, 5,676,333, and 5,251,846, all issued to Rethorst and all entitled xe2x80x9cSupersonic Aircraft Shock Wave Energy Recovery Systemxe2x80x9d (collectively xe2x80x9cthe Rethorst patentsxe2x80x9d), describe a supersonic aircraft having a spike extending from the front of the aircraft and a forward ring on the fuselage for eliminating a sonic boom. The spike is said to direct the bow shock onto the manifold ring which recovers the shock energy and converts it to useful work. The spike is said to be extendable, but it does not include a complex surface contour, and it is not disclosed to include a number of telescopically collapsible sections. Instead, it is disclosed as a single cylindrical member which tapers to a point at its leading end.
U.S. Pat. No. 3,643,901, issued to Patapis, discloses a ducted spike for attachment to a blunt body operating at supersonic speed for the purpose of receiving and diffusing oncoming air to reduce pressure drag on and erosion of the blunt body.
U.S. Pat. No. 3,425,650, issued to Silva, discloses an apparatus which can be extended on a boom from the front of an aircraft to deflect air outwardly therefrom.
U.S. Pat. No. 3,655,147, issued to Preuss, covers a device attached to the lower forebody of an aircraft for the purpose of reflecting pressure disturbances caused by the aircraft""s flight in directions away from the ground.
Although some of the foregoing references are directed to sonic boom mitigation, none of them address the sonic boom signature shaping techniques of the present invention.
The present invention provides an improvement in aircraft design which is directed to mitigating the effects of sonic booms at ground level. An aircraft according to the present invention includes a spike which extends from the aircraft""s nose in a direction substantially parallel to the aircraft""s length to effectively lengthen the aircraft. A longer aircraft generally is expected to produce a sonic boom of lesser amplitude at ground level than a shorter aircraft of similar weight because the pressure disturbance is distributed over a greater length. Therefore, a sonic boom created by an aircraft according to the present invention is expected to be of lesser intensity than a sonic boom created by a conventional supersonic aircraft of similar weight and otherwise similar size.
The spike can include several sections of varying cross-sectional area. The foremost, or farthest upstream, section of the spike preferably has a cross-sectional area which is characteristically small compared to that of the aircraft""s full fuselage or fuselage forebody. Generally, subsequent (farther aft), downstream sections of the spike progressively increase in cross-sectional area. However, a particular downstream section could have a smaller cross-sectional area than one or more upstream sections.
The transitions between sections preferably occur through curved or generally conical transition surfaces. However, other transition region contours are possible, as well. The foremost portion of the spike preferably tapers to a tip at its leading end, also through a curved, conical, or other transition region.
In preferred embodiments, the spike can be retracted into the fuselage when sonic boom mitigation is not needed or desired. For example, it may be desirable to retract the spike into the fuselage when the aircraft is flying at subsonic speeds, flying at supersonic speeds-over areas where sonic boom mitigation is deemed unnecessary (such as over an ocean), or is on the ground (to facilitate taxiing and parking).
The spike can be a single member. However the spike preferably includes two or more sections which can be collapsed telescopically to facilitate retraction of the spike into the fuselage and further to facilitate adjustment of the spike""s overall length and the relative position of the foregoing transitions between sections of varying cross-sectional area. For example, in a preferred embodiment, the spike includes a substantially cylindrical center section (which is the foremost section of the spike when the spike is fully or partially extended) surrounded by one or more overlapping, collapsible, annular sections. In other embodiments, the several sections can have other regular or irregular cross-sectional shapes. In such alternate embodiments, the spike can be a single member or it can be configured as two or more collapsible sections in a manner similar to that described above.
When an aircraft embodying such a spike is flown at supersonic speed, the tip of the spike causes an initial shock wave to be formed. Because at least the foremost portion of the spike""s cross-section is characteristically smaller than that of the full fuselage or fuselage forebody, this initial shock is of substantially weaker strength than the initial shock that would be generated by the full fuselage or fuselage forebody of an otherwise similar aircraft not having a spike. Further weak shocks are caused by the cross-sectional area transitions between adjacent telescoping sections (or similar discontinuities in a one-piece spike""s contour), as discussed above.
The position and shape of the foregoing transition regions define the strength and position of the weak shock waves created thereby. The position and shape of these transition regions are selected to reduce coalescence of the weak shocks into a strong sonic boom at the ground. The optimum position and shape of these transition regions are functions of several variables and can be expected to vary from aircraft to aircraft, based on the particular aircraft""s overall configuration. For example, the optimum position and shape of the transition regions may depend on the aircraft""s overall length, weight, fineness ratio, wing placement, engine placement, empennage design, etc. In some embodiments of the present invention, the position of such transition regions relative to each other and/or the aircraft""s fuselage can be adjusted on demand by incrementally extending or retracting particular sections of the spike.
A spike according to the present invention can be used in connection with conventional fuselage designs. It also can be used in connection with other fuselage designs, for example, without limitation, a fuselage configuration wherein the nose of the fuselage lies on a line substantially defining the bottom of the fuselage. When an aircraft embodying this design flies at supersonic speed, it creates an asymmetrical pressure distribution. The shock waves created by such an aircraft during normal supersonic flight propagate toward the ground with lesser intensity than in other directions. Detailed computational fluid dynamics (CFD) calculations and propagation analyses have shown that such an aircraft can be expected to produce a characteristically weaker acoustic signature at the ground than conventional aircraft. Thus, the foregoing fuselage shaping technique provides an important ingredient for shaping the sonic boom signature to permit supersonic flight over land. In alternate embodiments, at least the forward portion of the spike itself can be shaped in a manner similar to the novel fuselage discussed above. A spike embodying such a configuration causes the portions of the shock waves that propagate toward the ground to be of lesser intensity than the corresponding portions of the shock waves produced by an axisymmetric spike.
Similar benefits can be realized from the placement of a spike as described above at the rear of a supersonic aircraft. Accordingly, the present invention can be embodied as an aircraft having a spike projecting from the aft fuselage or empennage closure thereof in addition to or instead of the forward-projecting spike described above.